The lack of low cost access to orbit is the greatest single detriment to the near-term exploitation of space. The attempt to achieve orbital payload delivery for substantially less than $5,000/lb. (a representative figure for expended boosters) has yet to be realized. The attempt to lower the cost with the partially reusable NASA space shuttle was never realized primarily due to flaws in system architecture. This invention relates to the invention of a near-single-stage-to-orbit configuration specially tailored to deliver payloads efficiently to orbit and return the vehicle to earth while still providing an impressive payload weight to orbit and without resorting to large expendable propellant tankage or large staged boosters. Numerous attempts have been made since the first orbital flight (Sputnik) to provide designs that possess these capabilities. The most notable current activity is the National Aerospace Plane Project which would rely on technology advances in structural and heat shielding materials and air-breathing propulsion. For the horizontal takeoff mode specified the payload delivered to orbit is subject to question.
Numerous design studies for the pure rocket mode to orbit have been sponsored by the U.S. Government including horizontal and vertical ground-launch, air-launch, single-stage and dual-stage options. All have demonstrated insufficient cost leverage to move forward following the conceptual design. For the dual stage mode, the cost of concurrent development of two large hypersonic vehicles proved to be a major stumbling block. For the single-stage or near single-stage vehicle approach, the study results to date have been similarly disappointing. Vehicles are large with small payloads delivered to orbit.
The dilemma facing the reusable single-stage chemical rocket space-launch vehicle designer is illustrated in a relationship derived from the rocket equation:
.DELTA.V=g.times.Isp.times.log (W.sub.initial /W.sub.final)
Assuming .DELTA.V (including all losses)=29,000 feet/sec., vacuum Isp=460 sec. for a LOX-Hydrogen engine, and a fixed payload weight=10,000 lbs., the following relationship for the launch weight results: ##EQU1##
For a propellant to empty spacecraft weight ratio=6, there is no solution. For the ratio=7, the launch weight=490,000 lbs. Decreasing the ratio to 6.5 (a 7 percent change) nearly doubles the launch weight to 900,000 lbs. Since LOX-Hydrogen propellant has a density on the order of 20 lbs. /cu.ft., the space launch vehicle is in effect a flying-recoverable propellant tank making it extremely difficult to achieve a propellant to empty weight ratio in the range providing some assurance that low cost access to space can be achieved. Heretofore, configuration designs having horizontal landing capability have either employed wings (heavy) or are modifications of earlier tested shapes configured for a different application. None have made the propellant tankage an integral part of the lifting body aerodynamic design.